1. Field of the Invention
The present invention relates generally to gas turbine engines and, more particularly, to a clearance control assembly employing a one-piece continuous cylindrical housing for radially positioning shroud segments in response to temperature-controlled air flow in spiral passages defined between the housing and shroud segments.
2. Description of the Prior Art
Gas turbine engines typically include a core engine having a compressor for compressing air entering the core engine, a combustor where fuel is mixed with the compressed air and then burned to create a high energy gas stream, and a first or high pressure turbine which extracts energy from the gas stream to drive the compressor. In aircraft turbofan engines, a second turbine or low pressure turbine located downstream from the core engine extracts more energy from the gas stream for driving a fan. The fan provides the main propulsive thrust generated by the engine.
The engine components of the turbine and compressor includes a number of blades attached to a plurality of rotating rotor discs which are surrounded by a stationary shroud composed of axially arranged segments. In order to maintain engine efficiency, it is desirable to keep to a minimum the radial space, gap, or clearance between the tips of the rotor blades and the shroud segments. One of the major factors affecting efficiency of a gas turbine engine is this radial clearance between the adjacent rotating and non-rotating components.
If the radial clearance is too great, an unacceptable degree of gas leakage will occur with a resultant loss in efficiency. If the radial clearance is too little, there is a risk that under certain conditions contact will occur between the components.
The potential for contact occurring is particularly acute when the engine rotational speed is changing, either increasing or decreasing, since temperature differentials across the engine frequently result in the rotating and non-rotating components radially expanding and contracting at different rates. For instance, upon engine accelerations, thermal growth of the rotor typically lags behind that of the casing. During steady-state operation, the growth of the casing ordinarily matches more closely that of the rotor. Upon engine decelerations, the casing contracts more rapidly than the rotor.
Active control mechanisms, usually mechanically or thermally actuated, have been proposed in the prior art to maintain blade tip clearance substantially constant. Current turbofan engines have active control mechanisms only on high and low pressure turbines while the latter stages of the compressor, at best, have the rotor disks heated or cooled to control rotor blade-shroud clearances.
Typically, the engine compressor has an inner casing formed by two 180.degree. casing halves, or split casings, over its latter stages The split casings are attached together at horizontal flanges. During engine operation, the compressor is pressurized, causing the inner casing to assume a slightly oval configuration. The rotor blade-shroud clearances must be preset larger than desired in order to compensate for the oval configuration and avoid rubbing of the rotor blade against the shroud.
This split casing construction of the compressor inner casing with large preset compressor rotor blade-shroud clearances to avoid blade tip and shroud rubs results in reduced performance of and increased fuel consumption by the engine. Consequently, a need still exists for a construction for more effectively controlling clearances of rotor blade-shroud clearances.